This invention relates generally to gas turbine engines and more particularly to gas turbine engine components formed in part from high temperature foil materials.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed in turn by a turbine rotor.
Gas turbine engine hot section components, in particular the high pressure turbine section components, operate at extremely high temperatures and need to be cooled to have acceptable longevity. The tips of high pressure turbine (HPT) blades in particular are susceptible to high temperatures. The cooling is typically provided by extracting relatively cool air from an upstream location of the engine and routing the cooling air to components where it is needed. Conventionally the components to be cooled are hollow and have provisions for receiving and distributing the cooling air by various methods, for example the components may be film cooled by providing a plurality of passages which eject a blanket of cooling air over the surface of the component, or the components may be convectively cooled by causing the air to flow through various internal passages.
Conventional turbine blades often incorporate a squealer tip in which the sidewalls of the blade extend past the tip cap of the blade to define an open plenum therebetween. These extended tip walls are subject to severe oxidation and loss of material. There are two typical cooling designs for squealer tips. The first design uses impingement cooling from tip cap cooling holes. It is an effective cooling design. However, because of the angle required of the tip cap cooling holes, they cannot be formed in the completed airfoil and a braze-on tip cap having the cooling holes previously formed therein must be used. Unfortunately, a brazed joint has the drawback of increased susceptibility to oxidation compared to the base material of the blade. The second design is film cooling, which is used with blades having cast-in tip caps. Because the tip cap is cast-in, the tip cap cooling holes cannot be formed at the angle necessary to impingement cool the squealer tip. Therefore, film cooling on the outer surface of the pressure side wall is the main cooling source. However, the flow field near the tip region is quite complex and film cooling may not be fully satisfactory. It is further known to implement impingement cooling with a cast-in tip cap by forming a pair of coaxial angled holes through the squealer tip wall and the tip cap, and optionally, plugging the hole through the squealer tip wall. However, this still results in a blade having a tip constructed of a conventional superalloy, which is not as resistant to oxidation and material loss at high temperatures as desired.
Accordingly, there is a need for gas turbine engine airfoils having improved durability of the squealer tip walls.
The above-mentioned need is met by the present invention, which provides a turbine airfoil having an integral tip cap and a pair of tip walls which extend radially outward past the tip cap to form a squealer tip. A coaxial pair of first and second holes extend through the tip cap and one of the tip walls respectively at an acute inclination angle, with the first hole being disposed so as to direct impingement cooling air on the tip wall having the second hole. A protective layer comprising a high temperature foil is disposed on the tip wall having the second hole therein.